1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with a conical tip.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a compressor to compress air, a combustor to burn the compressed air with a fuel and produce a high temperature gas flow, and a turbine to convert the energy from the high temperature gas flow into mechanical energy used to drive the compressor and, in the case of an aero engine to drive a bypass fan, or in the case of an industrial gas turbine (IGT) engine to drive an electric generator.
The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the inlet temperature of the turbine is limited to the material properties of the first stage blades and vanes. Higher inlet turbine temperatures can be obtained by a combination of material properties (allowing for higher melting temperatures) and improved airfoil cooling. Since the compressed air used for airfoil cooling is bled off from the compressor, maximizing the amount of cooling while minimizing the amount of cooling air used is a major objective for the engine designer.
In a conical blade with cooling circuit, a serpentine tip turn will likely experience flow separation and recirculation issues. As a consequence of this, over temperatures occur at the locations of the blade tip turn regions corresponding to the flow separation. FIG. 1 shows a cut-away view of an aft flowing triple pass all convectively cooled turbine blade of the prior art. FIG. 2 shows a cross sectional view taken along the line A-A of the blade in FIG. 1. In the convectional cooling circuit of FIG. 1, the blade leading edge is cooled with a directed feed single pass radial flow channel. The leading edge cooling passage, in general, has a rough triangular shape as seen in FIG. 2 due to the narrowing of the airfoil wall at the leading edge. The inner surface area of the leading edge cooling passage reduces to the apex of an acute angle. The distribution of the cooling flow to the leading edge corner decreases and the substantial flow velocity as well as the internal heat transfer coefficient is reduced.
An alternative way to improve the airfoil leading edge cooling effectiveness while maintaining the same basic cooling circuit with the same amount of cooling flow is by the reduction of the airfoil leading edge cavity through flow area which increases the channel through flow velocity and therefore the resulting internal heat transfer coefficient. This is done by repositioning the leading edge rib forward as shown in FIG. 3. As a result of this modification, the blade tip turn cooling flow area ratio increases and yields a large unsupported mid-chord tip turn flow channel. The net impact due to this geometry change will enhance the blade tip turn flow separation and recirculation issues, especially for a blade with a conical tip design. As a consequence though, this design induces a higher blade tip turn loss and over temperature occurs at the location of the blade tip turn regions corresponding to the flow separation. This separation problem becomes even more pronounced for a blade with a conical tip. In addition, an increase of the airfoil mid-chord downward flowing channel flow area will reduce the through flow velocity and lower the internal heat transfer coefficient. Internal flow separation may occur for the mid-chord flow channel as well as the tip turn region when the internal Mach number is too low.